The present disclosure relates to a gas turbine engine and, more particularly, to a hollow blade having an internal damper.
Gas turbine engines, such as those that power modem commercial and military aircraft, include a fan section to propel the aircraft, compressor section to pressurize a supply of air from the fan section, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and generate thrust.
The gas turbine engines typically have rows of circumferentially spaced airfoils mounted on respective rotor disks for rotation about an engine axis. Advanced configurations feature shroudless hollow airfoils manufactured with lightweight materials. The airfoils are designed to high tolerances to accommodate significant operational requirements such as cross-winds and inlet distortion. These requirements result in airfoils that may be prone to high vibratory responses and possible aeroelastic instability within some operational speed ranges. To mitigate these effects, the airfoils may need to be damped.
One such damper is shown in U.S. Pat. No. 5,232,344, filed Jan. 17, 1992, where the damper operates under a centrifugal force that biases the damper against both a side skin at two transversely spaced locations of the airfoil and an airfoil face that faces radially inward. With more contemporary turbine engines, such as a geared turbine fan engine, slow fan running speeds are more common thus enabling production of aluminum alloy fan blades. Dampers, especially those that contact the aluminum blade skin at selected transverse locations, may be prone to causing unwanted wear on the blade itself.